text: At supersonic conditions, aircraft drag is originated due to:
  * Skin-friction drag due to shearing.
  * The wave drag due to the thickness (or volume) or zero-lift wave drag
  * Drag due to lift
  Therefore, the Drag coefficient on a supersonic airfoil is described by the following expression:
  C= C+ C+ C
  Experimental data allow us to reduce this expression to:
  C= C + KC</sub>
  Where C is the sum of C) and C, and k for supersonic flow is a function of the Mach number. The skin-friction component is derived from the presence of a viscous boundary layer which is infinitely close to the surface of the aircraft body. At the boundary wall, the normal component of velocity is zero; therefore an infinitesimal area exists where there is no slip. The zero-lift wave drag component can be obtained based on the supersonic area rule which tells us that the wave-drag of an aircraft in a steady supersonic flow is identical to the average of a series of equivalent bodies of revolution. The bodies of revolution are defined by the cuts through the aircraft made by the tangent to the fore Mach cone from a distant point of the aircraft at an azimuthal angle. This average is over all azimuthal angles. The drag due-to lift component is calculated using lift-analysis programs.  The wing design and the lift-analysis programs are separate lifting-surfaces methods that solve the direct or inverse problem of design and lift analysis.